In advanced turbine engines, combustion chamber temperatures may exceed 2400 degrees F. in order to increase the power output of the engine for high thrust and high efficiency. At the desired high temperature, the components in the hot engine section made of a Ni or Co based superalloy, even the advanced directionally solidified or single crystal nickel-based superalloys, will lose its load carrying function due to their mechanical strength being weakened due to overheating. The most advanced metallic coatings such as a McrAlY or platinum aluminized coating, which are applied onto the Ni or Co based superalloy turbine blades and vanes, are not adequate in protecting the superalloy article from oxidation since at this elevated temperature the metallic coating will oxidize at a very fast rate. In addition, these metallic coatings do not offer a sufficient thermal barrier to protect the superalloy article from overheating. The high operating temperatures have required the employment of thermal barrier coatings to protect both the base alloy and metallic coating from overheating, oxidation, erosion and corrosion etc. in the hostile operating environment of the engine which contains oxygen, abrasive particles and various contaminants.
Thermal barrier ceramic coatings are generally applied onto a bond coat that may be a simple aluminized or platinum aluminized coating or a MCrAlY coating, wherein M is Ni, Co and/or Fe, by air thermal spray (APS) or by electron beam physical vapor deposition (EBPVD). The most common thermal barrier coating material is a zirconia-based matrix with the addition of stabilizers such as yttria, calcia, magnesia, scandia, or halfnia. Some advanced thermal barrier coatings (“TBC”), consist of zirconia and/or halfnia based matrix with at least one dopant selected from rare earth oxides in the lanthanide group such as La, Ce, Pr, Nd, Pm, Sm, Eu, Gd, Tb, Dy, Ho, Er, Tm, or Yb, have been developed to achieve a higher thermal insulation with a lower thermal conductivity than the typical 7 wt. % yttria stabilized zirconia (7YSZ). The typical microstructure of an air plasma sprayed (“APS”) TBC predominantly is lamellar splats with anisotropic distribution of inter-splat lamellar pores that are mostly parallel to substrate and intra-splat cracks that are mostly perpendicular to substrate, as well as globular pores in various sizes. The typical microstructure of an electron beam physical vapor deposition (“EBPVD”) TBC is columnar grains with gaps between the columns and a feathery structure with intra-columnar micro pores. The columnar structure of EBPVD TBC allows the TBC to expand and contract without developing stress within itself during thermal cycling, which is the main reason that the EBPVD TBC exhibits a much longer lifetime than APS TBC.
Over the years, it has been recognized that further improvements to EBPVD TBC systems are needed to increase the lifetime of the coating system. Observations and failure investigations of overhauled engine blades and vanes have demonstrated that the TBC deterioration is primarily caused by CMAS and erosion attack. CMAS is a calcia-magnesia-alumina-silicate deposit originated from the ingestion of dust, sand, volcanic ashes, and runway debris with the intake of air in gas turbine engines. These CMAS elements deposit onto the surface of the TBC. At temperature in excess of 2100° F. during engine operation, these CMAS elements will melt and infiltrate into the gaps between the EBPVD TBC columns. When the engine shuts and cools down, the CMAS composites solidify into a fully dense domain and initiate large compressive stress within the TBC while it is contracting. When the compressive stress becomes too large, it causes a layer of the TBC to spall off at the depth of the CMAS penetration. Once the TBC is removed locally or entirely by CMAS attack, the bond coat and base alloy are subjected to high temperature-induced deterioration.
Erosion is a wear process that initiates when ingested particles such as sand, carbon from fuel combustion products, TBC coating particles from the combustion chamber and other foreign material impinge upon the TBC layer. The erodents in the high velocity gas stream strike the surface of turbine blades and vanes and remove the TBC. In the recent decades, low thermal conductivity TBCs (“low K TBC”) have been developed to meet the demand for increasing gas turbine operation temperatures. While lowering thermal conductivity of TBC is achieved either by modifying microstructure to incorporate more pores and cracks or by doping rare earth oxides into a zirconia matrix to form ternary or quaternary oxides, the erosion resistance of low K TBCs was ultimately sacrificed. To date, both EBPVD and APS low K TBCs have exhibited a higher erosion rate than the conventional 7YSZ TBC.
Efforts have been made to improve either erosion or CMAS resistance. U.S. Pat. No. 5,350,599 discloses an erosion resistant dense top layer produced by either halting or slowing the rotation while continuing deposition in order to make a densified layer. U.S. Pat. No. 5,683,825 discloses an erosion-resistant composition with Al2O3 or SiC dispersed within or overlaying the ceramic TBC layer. U.S. Pat. No. 5,714,202 discloses a diamond film deposited over columnar thermal barrier coatings for improving the erosion resistance. U.S. Pat. Nos. 5,792,521 and 6,054,184 disclose an alternative multilayer coating consists of a layer of 7YSZ for thermal insulation and a layer of Al2O3 for erosion resistance applied by sequentially deflecting an electronic beam at different ceramic ingots or use of a baffle to control vapor exposure. International Patent WO97/01436 discloses a continuous first sacrificial oxide or impermeable layer chosen from a metal oxide, metal carbide, metal nitride, metal silicate, or a precious metal on the outer surface of TBC combined with a top layer of non-wetting coating chosen from oxides, carbides, nitrides and silicates to resist CMAS infiltration and its reaction with the TBC.